launch vehicle avionics for passive thermal management · passive thermal management. william g....

Presented By William G. Anderson

Launch Vehicle Avionics for Passive Thermal Management William G. Anderson, Cameron Corday, Mike

DeChristopher, John Hartenstine, Taylor Maxwell, Carl Schwendeman, and Calin Tarau

Advanced Cooling Technologies, Inc.

Thermal & Fluids Analysis Workshop TFAWS 2014 August 4 - 8, 2014 NASA Glenn Research Center Cleveland, OH

TFAWS Active Thermal Paper Session

2 ADVANCED COOLING TECHNOLOGIES, INC.

ISO9001-2008 &AS9100-C Certified

Design Objectives

Develop a Launch Vehicle Avionics Passive Thermal Management System – Series of aluminum shelves with avionics boxes – Arbitrary location and size of the avionics boxes

Provide cooling for avionics – On ground prior to launch (indefinitely) – Purge duct – During launch (~10min) – Thermal storage – Avionics cooling on orbit

(indefinitely) – Radiator Heat Source

– 175 Watts per box (Max.) – 500 Watts per shelf (Max.) – 1-4 boxes per shelf



Pre-Launch Heat Sink

Heat Sink – Purge duct nitrogen sink: 67°F (~19.4°C) – Purge flow pressure drop cannot exceed 0.1 psid per shelf – Flow rates vary between 300 and 1800 scfm

– Distance between shelf and duct ranges from L = 12”-48”

– Duct diameter, D = 4”-10” – Baseplate allowable max:

+78°F (~25.5°C) – Baseplate allowable min:

+20°F (~6.7°C) – Purge duct is above all of the

shelves



Design Constraints

Temperatures

Baseplate Allowable Temperature Range -6.6 to 25.5°C Purge Duct Nitrogen Sink 19.4 °C

Purge Duct Aluminum, wall thickness: 1.6 mm (0.063 in.) Purge Duct Diameter 10 to 25 cm (4 to 10 in.) Purge Duct Length 20 to 96 cm (12 to 48 in.) Purge Duct Flow 8.5 to 51 m3/min (300 to 1800 cfm) Purge Flow Pressure Drop in Fin Stack 690 Pa (0.1 psid)/per shelf

Avionics Shelf Maximum power per shelf 500 W Shelf Length 152 cm (60 in.) Shelf Width 46 cm (18 in.) Shelf Thickness 3.8 cm (1.5 in.)

Avionics Boxes Maximum Power per box 175 W Maximum (Width x Depth x Height) 50 x 30 x 30 cm (20 x 12 x 12 in.) Average 30 x 30 x 30 cm (12 x 12 x 12 in.) Minimum 15 x 25 x 15 cm (6 x 10 x 6 in.)



Thermal Management System

The thermal management system has 5 components 1. A heat collection system to collect the heat from the electronics

boxes, and deliver it to a heat pipe 2. A heat transport system, consisting of riser heat pipes that transfer

heat from the avionics shelf to the heat sinks. – Thermosyphon on the ground – CCHP in space

3. A finned heat sink to reject the heat to the purge duct flow on the ground

4. A radiator to reject heat on orbit. – While CCHPs are discussed here, VCHPs would be necessary in

certain orbit to maintain the minimum electronics temperature 5. A thermal storage system, to accept the heat generated during

ascent



Thermal Management System Layout

Radiator Panel (w/ Embedded

Heat Pipes)

Finned HEX Assembly

CCHP Header

PCM (“Blue”)

Hi-K Plate (w/ Fully Embedded Heat Pipes) 60 in

18 in 12 – 48 in

48 in

92 in

20 in



Heat Collection System

Avionics Shelf requires – High effective thermal conductivity – Arbitrary location of avionics boxes and heat sources

Two possible solutions – Encapsulated pyrolytic graphite – High Conductivity (HiK™) plates with embedded heat pipes

Encapsulated pyrolytic graphite rejected

S. Kugler, “Aluminum Encapsulated APG High Conductivity Thermal Doubler” Spacecraft Thermal Control Workshop, El Segundo, CA, March 11-13, 2008

– Relatively low effective thermal conductivity, ~ 550 W/m K

– Thermal vias required due to the very low conductivity in the Z direction

– Location of vias is fixed during the fabrication process.

http://www.1-act.com/hik-plates-2/



Aluminum Plate HiK™ Plate

HiK™ plates have copper/water or copper/methanol heat pipes – Flatten, solder in machined slots – Can withstand thousands of freeze/thaw cycles – Operate up to 12 inches against gravity (if water is used) – Effective thermal conductivity of 500 – 1200 W/m K for terrestrial

applications, up to 2500 W/m K for spacecraft Identical Dimensions, 22°C Reduction in Peak Temperature

Measured

Embedded Heat Pipe Plates



Embedded Heat Pipe Plate Example

HiK™ plate with copper/water and methanol/water heat pipes was fabricated – Heat supplied by heaters at

one end of the plate – Removed by a 25°C cold plate

simulating a VCHP at the other end

– 127 cm (50 in.) by 61 cm (24 in.) by 0.79 cm (0.312 in.)

Test case: 95 W applied to heaters 1 and 4 – 50°C maximum allowable

Maximum plate temperature for this case was 43.9°C



HiK™ Plate: 95 W total at Heaters 1 and 4

Maximum plate temperature for this case was 43.9°C (18.9°C ΔT) Effective thermal conductivity of 2100 W/m K

TC 1 43.87TC 2 43.67TC 3 42.11TC 4 39.73TC 5 39.38TC 6 39.66TC 7 40.89TC 8 41.98TC 9 43.41TC 10 43.32TC 11 34.54TC 12 31.17TC 13 29.30TC 14 29.35TC 15 30.35TC 16 30.05TC 17 24.91TC 19 40.74TC 20 42.08TC 21 37.17TC 22 31.43TC 18 24.27

ambient 21.41power 95.10



Hi-K Plate Thermal Analysis

ΔT calculated between heat source and CCHP vapor

Hi-K Plate Module

Hi-K Plate Module Geometry • 18”x20”x0.25” • Cu/H2O heat pipes (51)

- OD = 0.25” - Wall thickness = 0.02” - Average Length = 11.25”

CCHP Headers



Results of Hi-K Plate Thermal Analysis

Worst Case: ΔT ~ 2°C Medium Case: ΔT ~ 1.4°C

Best Case: ΔT ~ 1.1°C

Thermal interface between flange and plate not considered



Heat Pipes/Heat Sinks

Standard grooved aluminum/ammonia heat pipes – Evaporator located underneath shelf – Allows electronics boxes to be mounted anywhere – 3 heat pipes based on transport capability

Finned Heat Sink

– Aluminum fins – Nitrogen Purge at 19.4°C (67°F) – ΔP < 230Pa (0.033 psid) per

header heat pipe Examine 300 and 1800 cfm

cases – Vary fin geometry – CFD simulations



CFD Example Geometry

– Nitrogen is flowing left to right – The heat pipe is the blue half circle (top picture) – The heat pipe fin is visible by velocity difference at

the fin boundaries – Results for 14 fin/in



300CFM CFD Results

Fin pitch iterations were carried out for a a fixed height of 10 in. – Higher fin pitch results in a higher pressure drop but lower temperature

difference (as expected) – 14 fin/in was selected as it meets the pressure drop requirement



Fin Configurations

To meet the 4.1°C ΔT and pressure drop requirements at 200W 300 CFM

– 25.4 cm (10 in.) tall heat pipe condenser – 14 fins/in. – Fin width 8.3 cm (3.25 in.) – Duct width 8.3 cm (3.25 in.) – Fin thickness 0.51 mm (0.02 in.)

1800 CFM – 26.7 cm (10.5 in.) tall heat pipe condenser – 12 fins/in. – Fin width 9.5 cm (3.75 in.) – Duct width 22.9 cm (9.0 in.) – Fin thickness 0.51 mm (0.02 in.)



Thermal Stack-Up (Pre-Launch)

ΔT = 2.0 °C

Tsink = 19.4 °C

Q = 175 W

ΔT ~ 1.0 °C

ΔT300CFM = 3.5 °C

Heat Source (175W)

Embedded Heat Pipes

PCM

CCHP Header

Finned HEX

Tvapor 1

Tvapor 2

Tplate

ΔT1800CFM = 3.1 °C



Thermal Storage During Launch

Thermal storage for 500W of heat from a shelf for ~10 minutes (i.e. 300kJ). – Hydride and PCM thermal storage were considered

Hydrides rejected, since don’t want to vent hydrogen – Rubitherm, (SP25A08) was selected as the PCM

Melts at 25°C – With this PCM, need to drop the purge temperature from 19.4 to 16.9°C,

to keep the PCM frozen before launch PCM was spread over half of the bottom of each one-third of the

shelf – PCM thickness of 2 mm (0.080 in.). – Conductivity enhanced with aluminum fins



Thermal Stack-Up (During Launch) - Worst Case -

ΔT = 1.7 °C

Plate Bottom

Plate Top

ΔTmelt,avg = 25 °C Q = 175 W

PCM ΔT based on steady-state analysis



Radiator Analysis

Heat pipe radiator Power is constant during mission at any value between 0 and 500W

– Emissivity…0.85 – View factor…0.9 – Minimum Efficiency…0.9 – Hottest sink…-40ºC – Coldest sink…-84.5ºC – Trans-lunar…-248.4ºC



Radiator Analysis

Radiator Geometry • Panel dimensions: 92”x48x”0.568” • 20 Embedded Al/NH3 heat pipes (0.5”) • 0.508” thick Al honey-comb structure • Two 0.03” Al face sheets



Radiator Analysis Shelf temperature as sink temperature changes is evaluated Shelf max. allowable temp. 25.5ºC Shelf max. allowable temp. -6.6ºC VCHP may be needed for certain orbits

Shelf Max. Allowable Temp.

Shelf Temp.

Shelf Min. Allowable Temp.



Thermal Stack-Up (During Orbit)

ΔT = 1.98 °C

ΔT ~ 1 °C

ΔT = 2 °C

Tsink = - 40°C

Tvapor 2

Tvapor 1

Tplate Q = 175 W

Tmax,surface



Summary of Thermal Stack-Up



System Mass/Shelf

Component Description Component Mass (kg) Quantity

Mass (kg)

Al Hi-K Plate (0.25") 7.38 1 7.38 Embedded Cu Pipes 0.06 153 9.18 Embedded Ti Pipes 0.02 153 3.06 Al/NH3 CCHP Header 1.09 3 3.27 Total PCM Package 0.93 3 2.80 Al Finned Stack (for duct) 300CFM 0.40 3 1.21 Al Finned Stack (for duct) 1800CFM 0.49 3 1.48 Radiator Panel 13.50 1 13.50 Embedded Al Pipes for Radiator NA 20 3.60 Total Mass w/ Cu Pipes (kg) 42.42 Total Mass w/ Ti Pipes (kg) 36.30



Experimental System

Low-cost, simplified version of the system was fabricated to verify the design during ground testing

One-third of a compete shelf Three components:

– High Conductivity Plate, or embedded heat pipe HiK™ plate. – Header heat pipe to purge duct – Removable Purge Flow Fin Heat Exchanger

Off-the-shelf heat sink



Experimental System Design Parameters

System designed for easy disassembly, to swap out components – TIM was Nusil CV2-2646 for all the make/break joints

Avionics HiK™ Shelf Heat Load 175 W Avionics Max. Heat Flux Density 0.51 W/cm2 (3.3 W/in2) Maximum Avionics Surface Temperature 26°C Avionics Shelf 50.8 x 45.7 cm (20 x 18 in.) Bolt hole pattern for avionics boxes 2 in. x 2 in. bolt hole pattern

Heat Sink Flow Rate 8.5 m3/min (300 scfm) Purge Flow Temperature 18 C Duct Diameter 10.2 cm (4 inch) Maximum pressure drop 690 Pa (0.1 psid) Heat Exchanger (Duct Length) Less than 30.5 cm (12 in.) Vertical Distance for Heat Transport 122 cm (48 in.)



HiK™ Plate Layout

36 embedded copper/water heat pipes Initial O.D. of 6.35 mm (0.25 in.) flattened to fit the plate Heat pipes laid out to avoid the mounting bolt holes Electrical heaters applied 175 W (0.51 W/cm2) over a 7.6 x 35.7 cm

(3 x 18 in.) strip at the edge of the plates (worst case).



Experimental Setup

Tests were conducted to verify basic functionality

TC 19

TC 21

TC 12

TC 22



Experimental Results – Basic Functionality Tests

Measured ΔT of roughly 15°C was above the predicted value of 13.5°C. – Much of the overall ΔT is due to the thermal interface materials (TIMs),

between the HiK™ plate and the CCHP, and the CCHP and the Heat Sink

– Can improve with a better thermal interface material, or by bonding the components together

High level goal is an avionics thermal management system with an overall thermal gradient of near 6°C. – Low-cost proof of concept system developed under this program

demonstrated that the components are conducive to a 6°C system

Test 1 Test 2 HiK™ Plate to Heat Pipe 5.87°C 5.58°C Heat Pipe Length 2.69°C 2.56°C Heat Pipe to Air 6.58°C 6.34°C Overall, Plate Surface to Air 15.13°C 14.40°C



Conclusions

A passive thermal management system was designed to cool avionics on a launch vehicle for 3 different thermal modes – Cooling for avionics while on ground prior to launch (indefinitely),

dumping the heat to flow in a purge duct. – Cooling for avionics during launch (~10min), storing the heat in a PCM

module – Cooling on-orbit (indefinitely), with a radiator.

The system has the following components – High Conductivity (HiK™) Plate with embedded heat pipes, where the

avionics are mounted – Three Constant Conductance Heat Pipe (CCHP) Header Heat Pipes,

that transfer heat to the purge duct of the radiator – Finned Heat Exchangers, the designated heat sink during pre-launch. – PCM modules, the designated heat sink during ascent. – Radiator, the designated heat sink on orbit.



Conclusions

Low-cost, simplified version of the system was designed and fabricated to verify the design during ground testing. – One-third of complete shelf 1. HiK™ plate 2. Al/ammonia heat pipe from the plate to the simulated duct 3. Low-cost, off the shelf heat sink in the duct

Functionality Testing – Measured ΔT of roughly 15°C – Above the predicted value of 13.5°C, due to the thermal interface

material – Overall thermal performance can be improved with a better interface

material Future Work

– System currently being tested at NASA Marshall



Acknowledgements

This project was sponsored by NASA Marshall Space Flight Center under Purchase Order Numbers 00096797 and NNM13AE60P – The Technical Monitor was Dr. Jeffery T. Farmer

Presented By William G. Anderson

[email protected]

Launch Vehicle Avionics for Passive Thermal Management William G. Anderson, Cameron Corday, Mike

DeChristopher, John Hartenstine, Taylor Maxwell, Carl Schwendeman, and Calin Tarau

Advanced Cooling Technologies, Inc.

Thermal & Fluids Analysis Workshop TFAWS 2014 August 4 - 8, 2014 NASA Glenn Research Center Cleveland, OH

TFAWS Active Thermal Paper Session

mailto:[email protected]

launch vehicle avionics for passive thermal management · passive thermal management. william g....

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